This application relates to a gas turbine engine, wherein a core engine is mounted separately from a propulsion unit.
Gas turbine engines are known, and have typically included a fan delivering a portion of air into a bypass duct, and a second portion of air into a core flow leading into a compressor section. The air is compressed in the compressor and delivered downstream into a combustion section where it is mixed with fuel and ignited. Products of this combustion pass across turbine rotors which are driven to rotate, and in turn rotate the compressor and fan section. Historically one turbine section drove both a compressor stage and a fan at the same speed. More recently it has been proposed to incorporate a gear reduction such as the fan can rotate at slower speeds than the compressor stage. With this arrangement, the outer diameter of the fan can increase, and the outer diameter of the turbine and compressor sections can decrease.
Historically, the fan and compressors have been mounted coaxially, and have been driven by turbines that are at a rear end of the engine, with the fan and compressor at a forward end. It has typically not been possible to service any portion of the engine, without removing the concentrically rotating turbines, compressors and fan as a combined unit. At a minimum, service is made complex by the inter-relationships of these sections.
Another challenge with mounting gas turbine engines relates to the so called “disk burst zone.” This zone is an area where broken pieces from a core engine could be driven.
The disk burst zone extends for approximately 30° about the last stage of the gas turbine engine. The gas turbine engine is typically mounted to an aircraft wing through a pylon. The aircraft wing also includes a fuel tank. There is a limitation on the mounting of current gas turbine engines in that the disk burst zone cannot extend through the fuel tank. Thus, gas turbine engines have typically been necessarily been mounted somewhat forwardly on the aircraft wing.